The present invention relates to a gas turbine engine combustor system, and more particularly, the invention relates to a fuel nozzle having a sensor to provide feedback control for the combustor system.
Gas turbine engines typically include a combustor system having multiple fuel nozzles. It is desirable to balance the fuel delivered by the fuel nozzles relative to one another to achieve uniform combustion across the combustor system and improve fuel efficiency. To this end, feedback control systems have been proposed and used on industrial gas turbine engines. Industrial gas turbine engines that use feedback controls employ a separate sensor, valve and fuel nozzle. Such a solution is too heavy for an aircraft gas turbine engine. Furthermore, separate components require additional penetrations into a housing diffuser case in the combustor system. Industrial gas turbine engine feedback control systems also have not been employed in gas turbine engines used on aircraft due to elevated temperatures and pressures in the combustor system of these engines.
What is needed is a feedback control system for a combustor system of an aircraft gas turbine engine that does not significantly increase weight, reduce reliability, or compromise the structural integrity of the engine. It is also desirable that the feedback control system be robust and easily replaced.